Propulsion system for an aircraft

ABSTRACT

A propulsion system for an aircraft includes a turbomachine, a primary fan, and an electric machine. The turbomachine includes a first turbine and a second turbine, with at least one of the first turbine or second turbine operably connected to the electric machine and the second turbine driving the primary fan. The propulsion system additionally includes an auxiliary propulsor assembly configured to be mounted at a location away from the turbomachine and the primary fan. The electric machine is in electrical communication with the auxiliary propulsor assembly for transferring power with the auxiliary propulsor assembly during operation of the propulsion system.

FIELD

The present subject matter relates generally to a propulsion system foran aircraft, and an aircraft including the same.

BACKGROUND

A conventional aircraft generally includes a fuselage, a pair of wings,and a propulsion system that provides thrust. The propulsion systemtypically includes at least two aircraft engines, such as turbofan jetengines. Each turbofan jet engine is mounted to a respective one of thewings of the aircraft, such as in a suspended position beneath the wing.

Additionally, turbofan jet engines are typically designed to maintain amaximum internal operating temperature below a certain threshold whenoperated at ground-level conditions (i.e., when ingesting air atambient, ground-level temperatures). However, once an aircraft reachescruise altitudes, the turbofan jet engines are ingesting air attemperatures much lower than ambient ground-level temperatures.Accordingly, with at least certain turbofan jet engines, there is roomto increase an internal operating temperature, and thus to increase anoverall pressure ratio of the engine, when operating at such cruisealtitudes.

Accordingly, a propulsion system that may more fully utilize anoperability range of a gas turbine engine during cruise operating modeswould be useful.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present disclosure, a propulsionsystem for an aircraft is provided. The propulsion system includes a gasturbine engine including a turbomachine and a primary fan. Theturbomachine includes a first turbine and a second turbine, and theprimary fan is driven by the second turbine. The propulsion system alsoincludes an electric machine operable with at least one of the firstturbine or the second turbine. The propulsion system additionallyincludes an auxiliary propulsor assembly configured to be mounted at alocation away from the gas turbine engine. The electric machine is inelectrical communication with the auxiliary propulsor assembly fortransferring power with the auxiliary propulsor assembly.

In another exemplary embodiment of the present disclosure an aircraft isprovided. The aircraft includes a first propulsion system including agas turbine engine having a primary fan and a turbomachine, an electricgenerator and an auxiliary propulsor assembly. The turbomachine isdrivingly connected to the electric generator, and the electricgenerator is electrically coupled to the auxiliary propulsor assemblyfor driving the auxiliary propulsor assembly. The aircraft additionallyincludes a second propulsion system including a gas turbine enginehaving a primary fan and a turbomachine, an electric generator and anauxiliary propulsor assembly. The turbomachine is drivingly connected tothe electric generator, and the electric generator is electricallycoupled to the auxiliary propulsor assembly for driving the auxiliarypropulsor assembly.

In an exemplary aspect of the present disclosure a method for operatinga propulsion system for an aircraft is provided. The propulsion systemincludes a gas turbine engine, an electric generator and an auxiliarypropulsor assembly, the gas turbine engine drivingly connected to theelectric generator and the electric generator electrically coupled tothe auxiliary propulsor assembly for driving the auxiliary propulsorassembly. The method includes operating the gas turbine engine in atakeoff operating mode such that a turbomachine of the gas turbineengine defines a first overall pressure ratio and provides the auxiliarypropulsor assembly with a first amount of electric power through theelectric generator. The method also includes operating the gas turbineengine in a cruise operating mode such that the turbomachine of the gasturbine engine defines a second overall pressure ratio and provides theauxiliary propulsor assembly with a second amount of electric powerthrough the electric generator. The second overall pressure ratio isgreater than the first overall pressure ratio.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a top view of an aircraft according to various exemplaryembodiments of the present disclosure.

FIG. 2 is a schematic, cross-sectional view of a gas turbine engine inaccordance with an exemplary embodiment of the present disclosure, whichmay be mounted to the exemplary aircraft of FIG. 1.

FIG. 3 is a schematic, cross-sectional view of an auxiliary propulsorassembly in accordance with an exemplary embodiment of the presentdisclosure, which may be mounted to the exemplary aircraft of FIG. 1.

FIG. 4 is a schematic, cross-sectional view of a gas turbine engine inaccordance with another exemplary embodiment of the present disclosure.

FIG. 5 is a schematic, cross-sectional view of an auxiliary propulsorassembly in accordance with another exemplary embodiment of the presentdisclosure.

FIG. 6 is a top view of an aircraft in accordance with another exemplaryembodiment of the present disclosure.

FIG. 7 is a top view of an aircraft in accordance with yet anotherexemplary embodiment of the present disclosure.

FIG. 8 is a top view of an aircraft in accordance with still anotherexemplary embodiment of the present disclosure.

FIG. 9 is a flow diagram of a method for operating a propulsion systemof an aircraft.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention. As used herein, theterms “first”, “second”, and “third” may be used interchangeably todistinguish one component from another and are not intended to signifylocation or importance of the individual components.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 provides a top view of anexemplary aircraft 10 as may incorporate various embodiments of thepresent disclosure. As shown in FIG. 1, the aircraft 10 defines alongitudinal centerline 14 that extends therethrough, a lateraldirection L, a forward end 16, and an aft end 18. Moreover, the aircraft10 includes a fuselage 12, extending longitudinally from the forward end16 of the aircraft 10 to the aft end 18 of the aircraft 10, and a wingassembly including a port side and a starboard side. More specifically,the port side of the wing assembly is a first, port side wing 20, andthe starboard side of the wing assembly is a second, starboard side wing22. The first and second wings 20, 22 each extend laterally outward withrespect to the longitudinal centerline 14. The first wing 20 and aportion of the fuselage 12 together define a first side 24 of theaircraft 10, and the second wing 22 and another portion of the fuselage12 together define a second side 26 of the aircraft 10. For theembodiment depicted, the first side 24 of the aircraft 10 is configuredas the port side of the aircraft 10, and the second side 26 of theaircraft 10 is configured as the starboard side of the aircraft 10.

Each of the wings 20, 22 for the exemplary embodiment depicted includesone or more leading edge flaps 28 and one or more trailing edge flaps30. The aircraft 10 further includes a vertical stabilizer 32 having arudder flap (not shown) for yaw control, and a pair of horizontalstabilizers 34, each having an elevator flap 36 for pitch control. Thefuselage 12 additionally includes an outer surface or skin 38. It shouldbe appreciated however, that in other exemplary embodiments of thepresent disclosure, the aircraft 10 may additionally or alternativelyinclude any other suitable configuration. For example, in otherembodiments, the aircraft 10 may include any other configuration ofstabilizer.

Referring now also to FIGS. 2 and 3, the exemplary aircraft 10 of FIG. 1additionally includes a propulsion system 50 having a first propulsorassembly 52 and a second propulsor assembly 54. FIG. 2 provides aschematic, cross-sectional view of the first propulsor assembly 52, andFIG. 3 provides a schematic, cross-sectional view of the secondpropulsor assembly 54. As is depicted, each of the first propulsorassembly 52 and second propulsor assembly 54 are configured asunder-wing mounted propulsor assemblies.

Referring particularly to FIGS. 1 and 2, the first propulsor assembly 52is mounted, or configured to be mounted, to the first side 24 of theaircraft 10, or more particularly, to the first wing 20 of the aircraft10. The first propulsor assembly 52 generally includes a turbomachineand a primary fan. More specifically, for the embodiment depicted thefirst propulsor assembly 52 is configured as a turbofan engine 100, withthe turbomachine being configured as a core turbine engine 102 and theprimary fan being configured as a fan 104 operable with the core turbineengine 102.

As shown in FIG. 2, the turbofan 100 defines an axial direction A₁(extending parallel to a longitudinal centerline 101 provided forreference) and a radial direction R₁. As stated, the turbofan 100includes the fan 104 and the core turbine engine 102 disposed downstreamfrom the fan 104.

The exemplary core turbine engine 102 depicted generally includes asubstantially tubular outer casing 106 that defines an annular inlet108. The outer casing 106 encases, in serial flow relationship, acompressor section including a booster or low pressure (LP) compressor110 and a high pressure (HP) compressor 112; a combustion section 114; aturbine section including a first, high pressure (HP) turbine 116 and asecond, low pressure (LP) turbine 118; and a jet exhaust nozzle section120.

The exemplary core turbine engine 102 of the turbofan 100 additionallyincludes one or more shafts rotatable with at least a portion of theturbine section and, for the embodiment depicted, at least a portion ofthe compressor section. More particularly, for the embodiment depicted,the turbofan 100 includes a high pressure (HP) shaft or spool 122, whichdrivingly connects the HP turbine 116 to the HP compressor 112.Additionally, the exemplary turbofan 100 includes a low pressure (LP)shaft or spool 124, which drivingly connects the LP turbine 118 to theLP compressor 110.

As stated, the primary fan of the first propulsor assembly 52 isconfigured as the fan 104 for the embodiment depicted. Further, theexemplary fan 104 depicted is configured as a variable pitch fan havinga plurality of fan blades 128 coupled to a disk 130 in a spaced apartmanner. The fan blades 128 extend outwardly from disk 130 generallyalong the radial direction R₁. Each fan blade 128 is rotatable relativeto the disk 130 about a respective pitch axis P by virtue of the fanblades 128 being operatively coupled to a suitable actuation member 132configured to collectively vary the pitch of the fan blades 128. The fan104 is mechanically coupled to the LP shaft 124. More particularly, thefan 104, including the fan blades 128, disk 130, and actuation member132, is mechanically coupled to the LP shaft 124 through a power gearbox134, and is rotatable about the longitudinal axis 101 by the LP shaft124 across the power gear box 134. The power gear box 134 includes aplurality of gears for stepping down the rotational speed of the LPshaft 124 to a more efficient rotational fan speed. Accordingly, the fan104 is powered by an LP system (including the LP turbine 118) of thecore turbine engine 102.

Referring still to the exemplary embodiment of FIG. 2, the disk 130 iscovered by rotatable front hub 136 aerodynamically contoured to promotean airflow through the plurality of fan blades 128. Additionally, theturbofan 100 includes an annular fan casing or outer nacelle 138 thatcircumferentially surrounds the fan 104 and/or at least a portion of thecore turbine engine 102. Accordingly, the exemplary turbofan 100depicted may be referred to as a “ducted” turbofan engine. It should beappreciated that the nacelle 138 is configured to be supported relativeto the core turbine engine 102 by a plurality ofcircumferentially-spaced outlet guide vanes 140. Moreover, a downstreamsection 142 of the nacelle 138 extends over an outer portion of the coreturbine engine 102 so as to define a bypass airflow passage 144therebetween.

Referring still to FIG. 2, the propulsion system 50 additionallyincludes an electric machine, which for the embodiment depicted isconfigured as an electric generator 56. The electric generator 56 is,for the embodiment depicted, positioned within the core turbine engine102 of the turbofan engine 100 and is in mechanical communication withone of the shafts of the turbofan engine 100. More specifically, for theembodiment depicted, the electric generator is driven by the first, HPturbine 116 through the HP shaft 122. The electric generator 56 isconfigured to convert mechanical power of the HP shaft 122 to electricpower. Accordingly, the electric generator 56 is powered by an HP system(including the HP turbine 116) of the core turbine engine 102.

Referring still to FIGS. 1 and 2, the propulsion system 50 depictedadditionally includes an electrical power bus 58 to allow the electricgenerator 56 to be in electrical communication with one or more othercomponents of the propulsion system 50 and/or the aircraft 10. For theembodiment depicted, the electrical power bus 58 includes one or moreelectrical lines 60 connected to the electric generator 56, and for theembodiment depicted, extending through one or more of the outlet guidevanes 140. Additionally, the propulsion system 50 depicted furtherincludes one or more energy storage devices 55 (such as one or morebatteries or other electrical energy storage devices) electricallyconnected to the electrical power bus 58 for, e.g., providing electricalpower to the auxiliary propulsor assembly 54 and/or receiving electricalpower from the gas turbine engine/first propulsion system 52. In certainexemplary embodiments, the one or more energy storage devices 55 may bepositioned proximate the auxiliary propulsor assembly 54 for weightdistribution purposes. Inclusion of the one or more energy storagedevices 55 may provide performance gains, and may increase a propulsioncapability of the propulsion system 50 during, e.g., transientoperations. More specifically, the propulsion system 50 including one ormore energy storage devices 55 may be capable of responding more rapidlyto speed change demands.

It should be appreciated, however, that in other embodiments, theelectric generator 56 may be positioned in any other suitable locationwithin the core turbine engine 102, or elsewhere. For example, theelectric generator 56 may be, in other embodiments, mounted coaxiallywith the HP shaft 122 within the turbine section, or alternatively maybe offset from the HP turbine 122 and driven through a suitable geartrain. Additionally, or alternatively, the electric generator 56 may bedriven by both the LP system (e.g., the LP shaft 124) and the HP system(e.g., the HP shaft 122) via a dual drive system. For example, a gearassembly, such as an epicyclic gear assembly, may be provided to allowboth the LP shaft 124 and HP shaft 122 to drive the electric generator56. Additionally, or alternatively still, in various other exemplaryembodiments, the electric machine/electric generator 56 may instead beoperable with just the LP system. For example, referring briefly to FIG.4, a propulsion system 50 including a turbofan engine 100 and electricmachine/electric generator 56 in accordance with another exemplaryembodiment of the present disclosure is provided. The exemplary turbofanengine 100 and electric machine/electric generator 56 of FIG. 4 areconfigured in substantially the same manner as the exemplary turbofanengine 100 and electric machine/electric generator 56 of FIG. 2.However, for the embodiment of FIG. 4, the electric machine/electricgenerator 56 is instead operable with the second turbine, or rather theLP turbine 118 (i.e., the same turbine driving the fan 104 of theexemplary turbofan engine 100) via the LP shaft 124.

It should further be appreciated that the exemplary turbofan engine 100depicted in FIG. 2 may, in other exemplary embodiments, have any othersuitable configuration. For example, in other exemplary embodiments, thefan 104 may not be a variable pitch fan, and further, in other exemplaryembodiments, the LP shaft 124 may be directly mechanically coupled tothe fan 104 (i.e., the turbofan engine 100 may not include the gearbox134). Further, it should be appreciated, that in other exemplaryembodiments, the turbofan engine 100 may instead be configured as anyother suitable aircraft engine including a turbomachine mechanicallycoupled to a primary fan. For example, in other embodiments, theturbofan engine 100 may instead be configured as a turboprop engine(i.e., the primary fan may be configured as a propeller), an unductedturbofan engine (i.e., the gas turbine engine may not include the outernacelle 138), etc.

Referring now particularly to FIGS. 1 and 3, the exemplary propulsionsystem 50 additionally includes the second propulsor assembly 54positioned, or configured to be positioned, at a location spaced apartfrom the first propulsor assembly 52 (including, e.g., the turbomachineand the primary fan). More specifically, for the embodiment depicted,the second propulsor assembly 54 is mounted at a location away from thefirst propulsor assembly 52 along the lateral direction L such that theyingest different airstreams along the lateral direction L. However, inother embodiments, the first and second propulsor assemblies 52, 54 mayeach be mounted to the aircraft 10 using a common mount. With such aconfiguration, however, the first and second propulsor assemblies 52, 54may still be positioned on the mount in a manner such that they arespaced apart from one another, e.g., along the lateral direction L suchthat they ingest different airstreams along the lateral direction L.Referring still to the exemplary embodiment of FIGS. 1 and 3, the secondpropulsor assembly 54 is mounted to one of the first side 24 or secondside 26 of the aircraft 10, e.g., to one of the first wing 20 or thesecond wing 22 of the aircraft 10. Notably, for the embodiment depictedin FIG. 1, the second propulsor assembly 54 is mounted to the secondside 26 of the aircraft 10, or rather to the second wing 22 of theaircraft 10.

Referring particularly to FIG. 3, the second propulsor assembly 54 isgenerally configured as an auxiliary propulsor assembly 200, defining anaxial direction A₂ extending along a longitudinal centerline axis 202that extends therethrough for reference, as well as a radial directionR₂. Additionally, the auxiliary propulsor assembly 200 generallyincludes an auxiliary fan 204 and an electric machine, which for theembodiment depicted is configured as an electric motor 206. For theembodiment depicted, the auxiliary fan 204 is rotatable about thecenterline axis 202. The auxiliary fan 204 includes a plurality of fanblades 208 and a fan shaft 210. The plurality of fan blades 208 areattached to/rotatable with the fan shaft 210 and spaced generally alonga circumferential direction of the auxiliary propulsor assembly 200 (notshown).

In certain exemplary embodiments, the plurality of fan blades 208 may beattached in a fixed manner to the fan shaft 210, or alternatively, theplurality of fan blades 208 may be rotatable relative to the fan shaft210, such as in the embodiment depicted. For example, the plurality offan blades 208 each define a respective pitch axis P2, and are attachedto the fan shaft 210 such that a pitch of each of the plurality of fanblades 208 may be changed, e.g., in unison, by a pitch change mechanism211. Changing the pitch of the plurality of fan blades 208 may increasean efficiency of the second propulsor assembly 54 and/or may allow thesecond propulsor assembly 54 to achieve a desired thrust profile. Withsuch an exemplary embodiment, the fan 204 may be referred to as avariable pitch fan.

Moreover, for the embodiment depicted, the auxiliary propulsor assembly200 depicted additionally include a fan casing or outer nacelle 212,attached to a core 214 of the auxiliary propulsor assembly 200 throughone or more struts or outlet guide vanes 216. For the embodimentdepicted, the outer nacelle 212 substantially completely surrounds thefan 204, and particularly the plurality of fan blades 208. Accordingly,for the embodiment depicted, the auxiliary propulsor assembly 200 may bereferred to as a ducted electric fan assembly.

Notably, the fan 204 of the auxiliary propulsor assembly 200 may definea fan pressure ratio. The fan pressure ratio may generally refer to aratio of a fan discharge pressure to a fan inlet pressure. As will bedescribed in greater detail below, the propulsion system 50 may beoperated during certain operations such that the auxiliary propulsorassembly provides relatively efficient thrust. For example, during,e.g., cruise operations of the propulsion system 50, the fan 204 of theauxiliary propulsor assembly 200 may define a fan pressure ratio of lessthan about 1.4:1. More specifically, in certain exemplary embodiments,the auxiliary fan 204 of the auxiliary propulsor assembly 200 maydefine, during cruise operations, a fan pressure ratio of less thanabout 1.3:1, such as less than about 1.2:1. It should be appreciated,that as used herein, terms of approximation, such as “about” or“approximately,” refer to being within a 10% margin of error.Additionally, the term “cruise operations” generally refers to a levelflight segment that occurs between an ascent phase and a descent phaseof the flight, at which the aircraft is designed for optimumperformance.

Referring still particularly to FIG. 3, the fan shaft 210 ismechanically coupled to the electric motor 206 within the core 214, suchthat the electric motor 206 drives the auxiliary fan 204 through the fanshaft 210. For the embodiment depicted, the electric motor 206 isconfigured as a variable speed electric motor, such that the electricmotor 206 may drive the auxiliary fan 204 at various rotational speedsdespite an amount of power provided thereto. Additionally, for theembodiment depicted, the auxiliary propulsor assembly 200 additionallyincludes an auxiliary propulsor gearbox 215 allowing for the rotationalspeed of the fan shaft 210 to be further increased or decreased relativeto a rotational speed of the electric motor 206. Accordingly, for theembodiment depicted, the electric motor 206 further drives the auxiliaryfan 204 across the auxiliary propulsor gearbox 215 and through the fanshaft 210.

Notably, however, in certain exemplary embodiments, the electric motor206 may be configured as a motor/generator. Accordingly, during, e.g.,emergency operations, the auxiliary propulsor assembly 200 may operateas a ram air turbine, such that inlet air to the auxiliary propulsorassembly 200 rotates the plurality of fan blades 208 of the fan 204, inturn rotating the electric motor/generator, allowing the electricmotor/generator to operate as an electric generator providing electricalpower to the power bus 58. Notably, with such an exemplary embodiment,the electric generator 56 of the turbofan engine 100 of FIG. 2 insteadoperates as an electric motor configured to receive power from theauxiliary propulsor assembly 200 and drive the core turbine engine 102.Moreover, it should be appreciated, that in other exemplary embodiments,the electric generator 56 may additionally be operable as an electricmotor to receive energy from a ground (or other external) power sourcefor e.g., starting the turbofan engine 100, and/or from an energystorage device, such as a battery, within the turbofan engine 100 oraircraft 10 for powering the turbofan engine 100.

The fan shaft 210 is supported by one or more bearings 218, such as theone or more roller bearings, ball bearings, or any other suitablebearings. Additionally, the electric motor 206 may be an inrunnerelectric motor (i.e., including a rotor positioned radially inward of astator), or alternatively may be an outrunner electric motor (i.e.,including a stator positioned radially inward of a rotor). As brieflynoted above, the electric generator 56 of the propulsion system 50 is inelectrical communication with the auxiliary propulsor assembly 200 forpowering the auxiliary propulsor assembly 200. More particularly, theelectric motor 206 of the auxiliary propulsor assembly 200 is inelectrical communication with the electrical power bus 58, which for theembodiment depicted includes one or more electrical lines 60electrically connected to the electric motor 206. Accordingly, theelectric motor 206 is more particularly in electrical communication withthe electrical power bus 58 through one or more electrical lines 60 ofthe electrical power bus 58, and the electrical power bus 58 may deliverpower to the electric motor 206 for driving the electric motor 206, andin turn driving the fan 204. Notably, for the embodiment depicted, theelectrical power bus 58 further includes one or more electricaldisconnects 61, such that the electrical power bus 58 may isolate one ormore components in the event of an electrical failure of one or morecomponents. The one or more electrical disconnects 61 may be manuallyoperated, or alternatively, may be automatically triggered in the eventof an electrical failure.

Referring again briefly to FIG. 1, the propulsion system 50 depicted, orrather, the electric power bus 58 depicted, additionally includes anelectric controller 62. The exemplary electric generator 56 depicted isin electrical communication with the auxiliary propulsor assembly 200through the electric controller 62 of the electric power bus 58. Theelectric controller 62 may be operably connected to one or moreadditional controllers of the aircraft, for controlling an amount ofpower provided to the auxiliary propulsor assembly 200.

It should be appreciated, however, that in other embodiments, theauxiliary propulsor assembly 200 may have any other suitableconfiguration. For example, referring now to FIG. 5, an auxiliarypropulsor assembly 200 in accordance with another exemplary embodimentof the present disclosure is depicted. The exemplary auxiliary propulsorassembly 200 of FIG. 5 may be configured in substantially the samemanner as the exemplary auxiliary propulsor assembly 200 of FIG. 3, andaccordingly, the same or similar numbers may refer to the same orsimilar part.

For example, the exemplary auxiliary propulsor assembly 200 of FIG. 5generally includes an auxiliary fan 204 including a plurality of fanblades 208 spaced generally along a circumferential direction of theauxiliary propulsor assembly 200 (not shown). Additionally, theplurality of fan blades 208 are attached to a fan shaft 210, with thefan shaft 210 supported by a one or more bearings 218. Each of theplurality of fan blades 208 includes an outer tip 220 along the radialdirection R2.

Additionally, an outer nacelle 212 and an electric motor 206 areprovided, with the outer nacelle 212 surrounding the auxiliary fan 204.However, for the embodiment depicted, the electric motor 206 is notconfigured to drive the auxiliary fan 204 through fan shaft 210.Instead, the electric motor 206 is at least partially integrated intothe tips 220 of one or more of the plurality of fan blades 208 forrotating the plurality of fan blades 208 directly. More specifically,the exemplary electric motor 206 of FIG. 5 generally includes a rotor222 and a stator 224. The rotor 222 is integrated into the tips 220 ofone or more of the plurality of fan blades 208 and the stator 224 ispositioned at least partially within the outer nacelle 212 of theauxiliary propulsor assembly 200. Notably, inclusion of an electricmotor 206 having such a configuration may allow for an electric motor206 having a reduced weight, which may in turn provide for additionalefficiency benefits of the auxiliary propulsion assembly 200.

A propulsion system in accordance with an exemplary embodiment of thepresent disclosure may provide for more efficient propulsion for anaircraft. For example, typically an overall pressure ratio of a gasturbine engine is limited by a temperature limit certain components ofthe gas turbine engine may withstand during takeoff operations (i.e., atsea level and typical sea level ambient temperatures). However, once anaircraft incorporating such a gas turbine engine reaches cruisingaltitudes, the ambient temperature of air ingested by the gas turbineengine is greatly reduced. Accordingly, the temperatures within such agas turbine engine are also greatly reduced. When the gas turbine engineis operated during cruise operations to simply provide thrust directlyand through, e.g., a fan (such as in a turbofan configuration), the gasturbine engine may not be utilizing its full potential. However, withthe present disclosure, energy may be extracted in certain exemplaryaspects from both an LP system (through the fan) and an HP system(through an electric generator) of the gas turbine engine, requiring theengine to operate at increased overall pressure ratios, and thusincreased temperatures. Specifically, extracting energy through both theLP system (i.e., through a primary fan) and the HP system (i.e., throughan electric generator) requires an increased amount of energy to begenerated by the turbomachine, which in turn requires an increasedoverall pressure ratio. Turbomachinery operating at an increased overallpressure ratio generally operates more efficiently. Notably, asdiscussed above, the energy extracted from the HP system may betransferred to an auxiliary fan of an auxiliary propulsor assembly togenerate additional thrust relatively efficiently.

Additionally, as used herein, “overall pressure ratio” of a gas turbineengine refers to a pressure ratio of a compressor section of aturbomachine of a gas turbine engine (e.g., for the embodiment of FIG. 2a ratio of a pressure immediately downstream of the HP compressor 112 toa pressure immediately upstream of the LP compressor). Similarly, an“overall pressure ratio” of a fan is a ratio of a pressure immediatelydownstream of the fan to a pressure immediately upstream of the fan.

Moreover, it should also be appreciated, that in still other exemplaryembodiments, the exemplary propulsion system 50 described above withreference to FIGS. 1 through 4 may be configured in any other suitablemanner. For example, in other exemplary embodiments, the fan 204 of theauxiliary propulsor assembly 200 may be mounted to a front hub of theauxiliary propulsor assembly 200. With such an embodiment, the fan shaft210 may be drivingly connected to the hub, or alternatively, theauxiliary propulsor assembly 200 may not include the fan shaft 210, andinstead the fan 204 and/or hub may be mounted directly on the electricmotor 206.

Additionally, in still other embodiments, the exemplary propulsionsystem may be integrated into an aircraft 10 in any other suitablemanner. For example, referring now to FIG. 6, an aircraft 10 inaccordance with another exemplary embodiment of the present disclosureis depicted. The exemplary aircraft 10 of FIG. 6 may be configured insubstantially the same manner as exemplary aircraft 10 of FIG. 1, andaccordingly, the same or similar numbers may refer to same or similarpart.

For example, the exemplary aircraft 10 of FIG. 6 generally includes afuselage 12 and a wing assembly, the wing assembly including a port sidewing 20 and a starboard side wing 22. Additionally, the exemplaryaircraft 10 of FIG. 6 includes a first propulsion system 250 inaccordance with an exemplary embodiment of the present disclosure. Thefirst propulsion system 250 may be configured in substantially the samemanner as exemplary propulsion system 50 described above with referenceto one or more of FIGS. 1 through 4. For example, the first propulsionsystem 250 may include a gas turbine engine 252, an electric generator(not shown), and an auxiliary propulsor assembly 254. The gas turbineengine 252 of the first propulsion system 250 may be drivingly connectedto the electric generator and the electric generator may be electricallycoupled to the auxiliary propulsor assembly 254 through a firstelectrical power bus 256 for driving the auxiliary propulsor assembly254. The gas turbine engine 252 of the first propulsion system 250 maygenerally include a primary fan and turbomachinery. For example, the gasturbine engine 252 may be configured as a turbofan engine (see FIG. 2).

However, for the embodiment depicted, the aircraft 10 further includes asecond propulsion system 258, the second propulsion system 258 alsoconfigured in accordance with an exemplary embodiment of the presentdisclosure. For example, the second propulsion system 258 may also beconfigured in substantially the same manner as exemplary propulsionsystem 50 described above with reference to one or more of FIGS. 1through 4. Specifically, the second propulsion system 258 may include agas turbine engine 260, an electric generator (not shown), and anauxiliary propulsor assembly 262. The gas turbine engine 260 of thesecond propulsion system 258 may be drivingly connected to the electricgenerator and the electric generator may be electrically coupled to theauxiliary propulsor assembly 262 through a second electrical power bus264 for driving the auxiliary propulsor assembly 262. The gas turbineengine 260 of the second propulsion system 258 may generally include aprimary fan and turbomachinery. For example, the gas turbine engine 260may be configured as a turbofan engine.

Specifically speaking, the gas turbine engine 252 of the firstpropulsion system 250 is mounted to the port side wing 20 and drives theauxiliary propulsor assembly 254 mounted to the starboard side wing 22.Additionally, the gas turbine engine 260 of the second propulsion system258 is mounted to the starboard side wing 22 and drives the auxiliarypropulsor assembly 262 mounted to the port side wing 20.

More generally, for the embodiment depicted, the gas turbine engine 252of the first propulsion system 250 is mounted to one of the port sidewing 20 or starboard side wing 22 and the auxiliary propulsor assembly254 of the first propulsion system 250 is mounted to the other of theport side wing 20 or starboard side wing 22. Similarly, the gas turbineengine 260 of the second propulsion system 258 is mounted to one of theport side wing 20 or the starboard side wing 22 and the auxiliarypropulsor assembly 262 of the second propulsion system 258 is mounted tothe other of the port side wing 20 or the starboard side wing 22.

Specifically, for the embodiment depicted, the gas turbine engine 252 ofthe first propulsion system 250 is mounted to the port side wing 20 andthe auxiliary propulsor assembly 254 of the first propulsion system 250is mounted to the starboard side wing 22. By contrast, the gas turbineengine 260 of the second propulsion system 258 is mounted to thestarboard side wing 22 and the auxiliary propulsor assembly 262 of thesecond propulsion system 258 is mounted to the port side wing 20.

An aircraft in accordance with the exemplary embodiment of FIG. 6 mayprovide for a more balanced aircraft from a thrust standpoint. Forexample, with such a configuration, if the first propulsion system 250fails, a propulsive force from both a port side 24 of the aircraft 10and from a starboard side 26 of the aircraft 10 is eliminated, and apropulsive force from both the port side 24 of the aircraft 10 and fromthe starboard side 26 of the aircraft 10 remains. The same is true ifthe second propulsion system 258 fails. Accordingly, elimination of oneof the first propulsion system 250 or second propulsion system 258 wouldnot result in an imbalanced propulsive force on the aircraft 10. Bycontrast, if the entire first propulsion system 250 were positioned onone side of the aircraft, and the entire second propulsion system 258were positioned on the other side of the aircraft, the verticalstabilizer 32 would need to be sized such that it could offset a momentgenerated by substantially all of the aircraft's thrust being generatedon one side of the aircraft. Notably, most current aircrafts having twopropulsion sources on opposing sides of the fuselage include arelatively large vertical stabilizer 32 precisely for this purpose.Accordingly, an aircraft configured in accordance with FIG. 6 (or FIG.7, discussed below) may allow for a reduction in size of the verticalstabilizer 32 of the aircraft 10, as in the event of a failure of one ofthe propulsion systems 250, 258 of the aircraft 10, the aircraft 10still receives propulsive force from both the port and starboard sides24, 26. Therefore, such a configuration may allow for a reduction in anamount of drag on the aircraft 10 during, e.g., cruising operations,which may result in an overall more efficient aircraft 10, providingpotentially significant cost benefits.

It should be appreciated, however, that in still other exemplaryembodiments, the exemplary propulsion system may be integrated into anaircraft 10 in any other suitable manner. For example, referring now toFIG. 7, an aircraft 10 in accordance with still another exemplaryembodiment of the present disclosure is depicted. The exemplary aircraft10 of FIG. 7 may be configured in substantially the same manner asexemplary aircraft 10 of FIG. 6, and accordingly, the same or similarnumbers may refer to same or similar part.

For example, the exemplary aircraft 10 of FIG. 7 generally includes afuselage 12 and a wing assembly, the wing assembly including a port sidewing 20 and a starboard side wing 22. Additionally, the exemplaryaircraft 10 of FIG. 6 includes a first propulsion system 250 inaccordance with an exemplary embodiment of the present disclosure. Forexample, the first propulsion system 250 includes a gas turbine engine252, an electric generator (not shown), and an auxiliary propulsorassembly 254. The gas turbine engine 252 of the first propulsion system250 may be drivingly connected to the electric generator and theelectric generator may be electrically coupled to the auxiliarypropulsor assembly 254 through a first electrical power bus 256 fordriving the auxiliary propulsor assembly 254. The gas turbine engine 252of the first propulsion system 250 may generally include a primary fanand turbomachinery. For example, the gas turbine engine 252 may beconfigured as a turbofan engine (see FIG. 2).

Additionally, the exemplary aircraft 10 of FIG. 7 includes a secondpropulsion system 258, the second propulsion system 258 also configuredin accordance with an exemplary embodiment of the present disclosure.For example, the second propulsion system 258 includes a gas turbineengine 260, an electric generator (not shown), and an auxiliarypropulsor assembly 262. The gas turbine engine 260 of the secondpropulsion system 258 may be drivingly connected to the electricgenerator and the electric generator may be electrically coupled to theauxiliary propulsor assembly 262 through a second electrical power bus264 for driving the auxiliary propulsor assembly 262. The gas turbineengine 260 of the second propulsion system 258 may generally include aprimary fan and turbomachinery. For example, the gas turbine engine 260may be configured as a turbofan engine (see FIG. 2).

Moreover, as with the embodiment of FIG. 6, the exemplary gas turbineengine 252 of the first propulsion system 250 is mounted to the portside wing 20 and drives the auxiliary propulsor assembly 254 mounted tothe starboard side wing 22. Additionally, the gas turbine engine 260 ofthe second propulsion system 258 is mounted to the starboard side wing22 and drives the auxiliary propulsor assembly 262 mounted to the portside wing 20.

Notably, referring back briefly to FIG. 6, for the embodiment of FIG. 6,the gas turbine engine 260 of the second propulsion system 258 ispositioned closer to the fuselage 12 than the auxiliary propulsorassembly 254 of the first propulsion system 250 on the starboard sidewing 22. Additionally, the auxiliary propulsor assembly 262 of thesecond propulsion system 258 is similarly positioned closer to thefuselage 12 than the gas turbine engine 252 of the first propulsionsystem 250 on the port side wing 20. Similarly, referring again to FIG.7, for the embodiment depicted, the gas turbine engine 260 of the secondpropulsion system 258 is positioned closer to the fuselage 12 than theauxiliary propulsor assembly 254 of the first propulsion system 250 onthe starboard side wing 22. By contrast, however, for the embodiment ofFIG. 7, the gas turbine engine 252 of the first propulsion system 250 ispositioned closer to the fuselage 12 than the auxiliary propulsorassembly 262 of the second propulsion system 258 on the port side wing20.

As will be appreciated, such a configuration may provide for a moreequally balanced aircraft 10 (from a weight standpoint). For example, asis depicted in FIG. 7, the gas turbine engine 252 of the firstpropulsion system 250 and the gas turbine engine 260 of the secondpropulsion system 258, while positioned on opposing sides of thefuselage 12, are spaced approximately the same distance from thecenterline 14 along the lateral direction L. Similarly, the auxiliarypropulsor assembly 254 of the first propulsion system 250 and theauxiliary propulsor assembly 262 of the second propulsion system 258,while also positioned on opposing sides of the fuselage 12, are alsospaced approximately the same distance from the centerline 14 along thelateral direction L. As it is possible that the gas turbine engines 252,260 will weigh more than the auxiliary propulsor assemblies 254, 260,the configuration of FIG. 7 may provide for a more equally balanced(weight-wise) aircraft 10.

It should be appreciated, however, that the exemplary propulsion systemsdescribed above, and the exemplary aircraft configurations incorporatingsuch propulsion systems described above, are provided for exemplarypurposes only. In other exemplary embodiments, the propulsion systemsand/or aircraft may have any other suitable configuration. For example,in other exemplary embodiments, a propulsion system in accordance withthe present disclosure may include a plurality of auxiliary propulsorassemblies mounted in any suitable configuration, each driven by anelectric generator of the propulsion system. Additionally, in stillother embodiments, one or more of the propulsion systems may includeenergy storage devices (such as batteries). Further, for example,referring now to FIG. 8, a propulsion system in accordance with stillanother exemplary embodiment is provided. The exemplary propulsionsystem of FIG. 8 may be configured in substantially the same manner asthe exemplary propulsion system described above with reference to FIG.6. According, the same or similar numbers may refer to same or similarpart.

For example, the exemplary aircraft 10 of FIG. 8 generally includes afuselage 12 and a wing assembly, the wing assembly including a port sidewing 20 and a starboard side wing 22. Additionally, the exemplaryaircraft 10 of FIG. 8 includes a first propulsion system 250 inaccordance with an exemplary embodiment of the present disclosure. Forexample, the first propulsion system 250 includes a gas turbine engine252, an electric generator (not shown), and an auxiliary propulsorassembly 254. Additionally, the exemplary aircraft 10 of FIG. 8 includesa second propulsion system 258, the second propulsion system 258 alsoconfigured in accordance with an exemplary embodiment of the presentdisclosure. For example, the second propulsion system 258 includes a gasturbine engine 260, an electric generator (not shown), and an auxiliarypropulsor assembly 262.

However, for the exemplary embodiment of FIG. 8, the first propulsionsystem 250, including the gas turbine engine 252 and auxiliary propulsorassembly 254, is mounted to the port side wing 20. Additionally, for theembodiment of FIG. 8, the second propulsion system 258, including thegas turbine engine 260 and the auxiliary propulsor assembly 262, ismounted to the starboard side wing 22. Notably, for the embodimentdepicted, the gas turbine engines 252, 260 of the first and secondpropulsion systems 250, 258 are each mounted closer to a longitudinalcenterline 14 of the aircraft 10 relative to the respective auxiliarypropulsor assemblies 254, 262 of the first and second propulsion systems250, 258. However, in other embodiments, the auxiliary propulsorassemblies 254, 262 may instead be mounted closer to the longitudinalcenterline 14 of the aircraft 10.

Additionally, in still other exemplary embodiments of the presentdisclosure, the propulsion systems, and the aircraft incorporating suchpropulsion systems, may be configured in still other suitable manners.For example, in still other exemplary embodiments, one or both of a gasturbine engine and an auxiliary propulsor assembly of a propulsionsystem may be mounted to a fuselage of the aircraft proximate a tail endof the aircraft. For example, the propulsion system may include the gasturbine engine mounted to one side of the fuselage and the auxiliarypropulsor assembly mounted to an opposite side of the fuselage (e.g.,the gas turbine engine mounted to one of a starboard side 26 or portside 24 of a fuselage 12 of an aircraft 10 proximate a tail end 18, andthe auxiliary propulsor assembly mounted to the other of the starboardside 26 or port side 24 of the fuselage 12 of the aircraft 10 proximatethe tail end 18). Moreover, in still other exemplary embodiments, one orboth of the gas turbine engine and the auxiliary propulsor assembly ofthe propulsion system may be mounted to a stabilizer of the aircraft,such as to a vertical stabilizer of the aircraft. In either of theseembodiments, the propulsion systems may further include one of the gasturbine engine or auxiliary propulsor assembly mounted in an under-wingconfiguration as well.

Referring now to FIG. 9, a flow diagram of a method (300) for operatinga propulsion system of an aircraft is provided. The exemplary method(300) may be utilized with one or more embodiments of the exemplarypropulsion systems and aircraft described above with reference to FIGS.1 through 8. For example, the exemplary method (300) may be utilizedwith a propulsion system including a gas turbine engine, an electricgenerator, and an auxiliary propulsor assembly. The gas turbine enginemay include a primary fan and a turbomachine, and further may drivinglybe connected to the electric generator. Additionally, the electricgenerator may be electrically coupled to the auxiliary propulsorassembly for driving the auxiliary propulsor assembly.

For the exemplary aspect depicted, the exemplary method (300) includesat (302) operating the gas turbine engine in a takeoff operating mode,such that the turbomachine defines a first overall pressure ratio andprovides the auxiliary propulsor assembly with a first amount ofelectric power from the electric generator. Additionally, at (304) themethod (300) includes operating the gas turbine engine in a cruiseoperating mode, such that the turbomachine defines a second overallpressure ratio and provides the auxiliary propulsor assembly with asecond amount of electric power from the electric generator.

For the exemplary aspect depicted, the second overall pressure ratiodefined by the turbomachine of the gas turbine engine at (304) isgreater than the first overall pressure ratio defined by theturbomachine of the gas turbine engine at (302). More specifically, inat least certain exemplary aspects, the second overall pressure ratiomay be at least about 5% greater, at least about 10% greater, or atleast about 20% greater than the first overall pressure ratio. Further,in at least certain exemplary aspects, the second amount of powerprovided to the auxiliary propulsor assembly at (304) may besubstantially equal to the first amount of power provided to theauxiliary propulsor assembly at (302). Accordingly, the method 300 mayinclude providing a substantially constant amount of power to theauxiliary propulsor assembly during takeoff and cruise. For example, incertain exemplary aspects, the method 300 may include providing asubstantially constant amount of power to the auxiliary propulsorthroughout a flight envelope.

As will be appreciated, an inlet air temperature and density for theturbomachine of the gas turbine engine during takeoff operatingconditions is generally higher than an inlet air temperature and densityfor the turbomachine of the gas turbine engine during cruise operatingconditions. Such allows for the turbomachine of the gas turbine engineto produce more power during takeoff as compared to during cruise.Accordingly, providing a substantially constant amount of power to theauxiliary propulsor during takeoff and cruise operating conditionsresults in a smaller fraction of power extraction at takeoff as comparedto the relatively larger fraction of power extraction at cruise. Thefraction of power extraction refers to a ratio of power provided to theauxiliary propulsor from the turbomachine to a total amount of powergenerated by the turbomachine.

The relatively larger fraction of power extraction at cruise isaccomplished, at least in part, by increasing a core speed of theturbomachine, and accordingly increasing an overall pressure ratio ofthe turbomachine. Notably, a decreased inlet air temperature allows forthe increase in overall pressure ratio of the turbomachine during cruiseoperations. For example, the turbomachine is typically operated to amaximum compressor discharge temperature and/or exhaust temperature. Byreducing an inlet air temperature, a greater pressure increase isallowed across a compressor section of the turbomachine, whilemaintaining the compressor discharge temperature and/or exhausttemperature at or below the maximum compressor discharge temperatureand/or exhaust temperature.

Further, it will be appreciated, that by increasing the core speed andoverall pressure ratio of the turbomachine during cruise operations, anincreased marginal fuel efficiency is accomplished for the turbomachineof the gas turbine engine. For example, the energy extracted by thegenerator of the propulsion system drives the auxiliary propulsor,effectively increasing an overall bypass ratio of the gas turbineengine, and thus increasing its overall propulsive efficiency, whilesimultaneously increasing the turbomachine efficiency due to the higheroverall core speed and overall pressure ratio of the turbomachine.

Additionally, as is depicted schematically in FIG. 9, in certainexemplary aspects of the present disclosure, the propulsion systemdescribed above with reference to (302) and (304) may be a firstpropulsion system and the aircraft may further include a secondpropulsion system. The second propulsion system may include a second gasturbine engine, a second electric generator, and a second auxiliarypropulsor assembly. The second gas turbine engine may include a secondprimary fan and second a turbomachine, and further may drivingly beconnected to the second electric generator. Additionally, the secondelectric generator may be electrically coupled to the second auxiliarypropulsor assembly for driving the second auxiliary propulsor assembly.

With such an exemplary aspect, the method (300) may further include at(306) operating the second gas turbine engine of the second propulsionsystem in a takeoff operating mode such that the second turbomachine ofthe second gas turbine engine defines a third overall pressure ratio andprovides the second auxiliary propulsor assembly of the secondpropulsion system with a third amount of electric power through thesecond electric generator of the second propulsion system. Moreover, themethod (300) may further include at (308) operating the second gasturbine engine in a cruise operating mode such that the secondturbomachine of the second gas turbine engine defines a fourth overallpressure ratio and provides the second auxiliary propulsor assembly witha fourth amount of electric power through the second electric generator.The fourth overall pressure ratio may be greater than the third overallpressure ratio (such as at least about 5% greater), and in certainexemplary aspects, the fourth amount of electric power may besubstantially equal to the third amount of electric power.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A propulsion system for an aircraft, thepropulsion system comprising: a gas turbine engine comprising aturbomachine and a primary fan, the turbomachine comprising a firstturbine and a second turbine, and the primary fan driven by the secondturbine; an electric machine operable with at least one of the firstturbine or the second turbine; and an auxiliary propulsor assemblyconfigured to be positioned at a location spaced apart from the gasturbine engine, the electric machine in electrical communication withthe auxiliary propulsor assembly for transferring power with theauxiliary propulsor assembly, wherein the propulsion system isconfigured to have a first fraction of power extraction when the gasturbine engine is in a takeoff operating mode and a second fraction ofpower extraction when the gas turbine engine is in a cruise operatingmode, the first and second fraction of power extraction each being aratio of power provided to the auxiliary propulsor assembly from theturbomachine to an overall amount of power generated by theturbomachine, wherein the first fraction of power extraction is lessthan the second fraction of power extraction.
 2. The propulsion systemof claim 1, wherein the electric machine is an electric generator drivenby the first turbine, and wherein the electric machine is in electricalcommunication with the auxiliary propulsor assembly for powering theauxiliary propulsor assembly.
 3. The propulsion system of claim 1,further comprising: an electric power bus, wherein the electric machineis in electrical communication with the auxiliary propulsor assemblythrough the electric power bus.
 4. The propulsion system of claim 3,wherein the electrical power bus comprises one or more electricaldisconnects for isolating electrical failures of the propulsion system.5. The propulsion system of claim 1, wherein the auxiliary propulsorassembly comprises an electric motor and an auxiliary fan, wherein theelectric motor is in electrical communication with the electric machineand drives the auxiliary fan.
 6. The propulsion system of claim 5,wherein the auxiliary propulsor assembly further comprises a nacellesurrounding the auxiliary fan, wherein the auxiliary fan comprises aplurality of fan blades having outer tips, wherein the electric motorcomprises a rotor and a stator, and wherein the rotor is attached to orintegrated into the outer tips of the plurality of fan blades.
 7. Thepropulsion system of claim 5, wherein the auxiliary propulsor assemblyfurther comprises a fan shaft, wherein the electric motor drives theauxiliary fan through the fan shaft.
 8. The propulsion system of claim5, wherein the auxiliary propulsor assembly further comprises a gearbox.9. The propulsion system of claim 5, wherein the electric motor is avariable speed motor.
 10. The propulsion system of claim 1, wherein theelectric machine is operable as an electric motor configured to receivepower from the auxiliary propulsor assembly.
 11. The propulsion systemof claim 1, wherein the auxiliary propulsor assembly comprises anauxiliary fan, and wherein the auxiliary fan is configured as a variablepitch fan.
 12. The propulsion system of claim 1, wherein the electricmachine is an electric generator driven by the second turbine.
 13. Thepropulsion system of claim 1, wherein the auxiliary propulsor assemblycomprises an auxiliary fan, and wherein the auxiliary fan defines a fanpressure ratio during cruise operations less than about 1.4:1.
 14. Anaircraft comprising: a first propulsion system comprising a gas turbineengine having a primary fan and a turbomachine, an electric generator,and an auxiliary propulsor assembly, the turbomachine drivinglyconnected to the electric generator, and the electric generatorelectrically coupled to the auxiliary propulsor assembly for driving theauxiliary propulsor assembly; and a second propulsion system comprisinga gas turbine engine having a primary fan and a turbomachine, anelectric generator, and an auxiliary propulsor assembly, theturbomachine drivingly connected to the electric generator, and theelectric generator electrically coupled to the auxiliary propulsorassembly for driving the auxiliary propulsor assembly, wherein theauxiliary propulsor assembly of the first propulsion system and theauxiliary propulsor assembly of the second propulsion system eachcomprise an auxiliary fan, wherein the first propulsion system isconfigured to have a first fraction of power extraction when the gasturbine engine is in a takeoff operating mode and a second fraction ofpower extraction when the gas turbine engine is in a cruise operatingmode, the first and second fraction of power extraction each being aratio of power provided to the auxiliary propulsor assembly from theturbomachine to an overall amount of power generated by theturbomachine, wherein the first fraction of power extraction is lessthan the second fraction of power extraction.
 15. The aircraft of claim14, further comprising: a wing assembly comprising a port side and astarboard side; wherein the gas turbine engine of the first propulsionsystem is mounted to one of the port side or starboard side of the wingassembly and wherein the auxiliary propulsor assembly of the firstpropulsion system is mounted to the other of the port side or starboardside of the wing assembly; and wherein the gas turbine engine of thesecond propulsion system is mounted to one of the port side or starboardside of the wing assembly and wherein the auxiliary propulsor assemblyof the second propulsion system is mounted to the other of the port sideor starboard side of the wing assembly.
 16. The aircraft of claim 15,wherein the aircraft defines a longitudinal centerline and a lateraldirection, wherein the gas turbine engine of the first propulsion systemis mounted to the port side and the gas turbine engine of the secondpropulsion system is mounted to the starboard side, wherein the gasturbine engines of the first and second propulsion systems are spacedsubstantially equally from the longitudinal centerline along the lateraldirection, wherein the auxiliary propulsor assembly of the firstpropulsion system is mounted to the starboard side and the auxiliarypropulsor assembly of the second propulsion system is mounted to theport side, wherein the auxiliary propulsor assemblies of the first andsecond propulsion systems are each positioned outward of the gas turbineengines and spaced substantially equally from the longitudinalcenterline along the lateral direction.
 17. The aircraft of claim 14,further comprising: a wing assembly comprising a port side and astarboard side; wherein the gas turbine engine and auxiliary propulsorassembly of the first propulsion system are each mounted to the portside of the wing assembly; and wherein the gas turbine engine andauxiliary propulsor assembly of the second propulsion system are eachmounted to the starboard side of the wing assembly.
 18. The aircraft ofclaim 14, wherein the auxiliary propulsor assembly of the firstpropulsion system and the auxiliary propulsor assembly of the secondpropulsion system each comprise an electric motor, wherein the electricmotors are each in electrical communication with the respective electricgenerators and drive the respective auxiliary fans.
 19. The aircraft ofclaim 14, wherein the auxiliary propulsor assembly comprises anauxiliary fan, and wherein the auxiliary fan defines a fan pressureratio during cruise operations less than about 1.4:1.
 20. A method foroperating a propulsion system for an aircraft, the propulsion systemcomprising a gas turbine engine, an electric generator, and an auxiliarypropulsor assembly, the gas turbine engine drivingly connected to theelectric generator and the electric generator electrically coupled tothe auxiliary propulsor assembly for driving the auxiliary propulsorassembly, the method comprising: operating the gas turbine engine in atakeoff operating mode such that a turbomachine of the gas turbineengine defines a first overall pressure ratio and provides the auxiliarypropulsor assembly with a first amount of electric power through theelectric generator; and operating the gas turbine engine in a cruiseoperating mode such that the turbomachine of the gas turbine enginedefines a second overall pressure ratio and provides the auxiliarypropulsor assembly with a second amount of electric power through theelectric generator, the second overall pressure ratio being greater thanthe first overall pressure ratio; wherein operating the gas turbineengine in the takeoff operating mode includes operating the propulsionsystem to define a first fraction of power extraction, wherein operatingthe gas turbine engine in the cruise operating mode includes operatingthe propulsion system to define a second fraction of power extraction,wherein the fraction of power extraction refers to a ratio of powerprovided to the auxiliary propulsor assembly from the turbomachine to anoverall amount of power generated by the turbomachine, and wherein thefirst fraction of power extraction is less than the second fraction ofpower extraction.
 21. The method of claim 20, wherein the auxiliarypropulsor assembly comprises an auxiliary fan, and wherein operating thegas turbine engine in a cruise operating mode includes operating thepropulsion system such that the auxiliary fan defines a fan pressureratio during cruise operations less than about 1.4:1.